The present invention pertains generally to electric plasma thrusters and more particularly to Hall field thrusters, which are sometimes called Hall accelerators.
The Hall plasma accelerator is an electrical discharge device in which a plasma jet is accelerated by a combined operation of axial electric and magnetic fields applied in a coaxial channel. The conventional Hall thruster overcomes the current limitation inherent in ion diodes by using neutralized plasma, while at the same time employing radial magnetic fields strong enough to inhibit the electron flow, but not the ion flow. Thus, the space charge limitation is overcome, but the electron current does draw power. Hall thrusters are about 50% efficient. Hall accelerators do provide high jet velocities, in the range of 10 km/s to 20 km/s, with larger current densities, about 0.1 A/cm2, than can conventional ion sources.
Hall plasma thrusters for satellite station keeping were developed, studied and evaluated extensively for xenon gas propellant and jet velocities in the range of about 15 km/s, which requires a discharge voltage of about 300 V. Hall thrusters have been developed for input power levels in the general range of 0.5 kW to 10 kW. While all Hall thrusters retain the same basic design, the specific details of an optimized design of Hall accelerators vary with the nominal operating parameters, such as the working gas, the gas flow rate and the discharge voltage. The design parameters subject to variation include the channel geometry, the material, and the magnetic field distribution.
A. V. Zharinov and Yu. S. Popov, xe2x80x9cAcceleration of plasma by a closed Hall currentxe2x80x9d, Sov. Phys. Tech. Phys. 12, 1967, pp. 208-211 describe ideas on ion acceleration in crossed electric and magnetic field, which date back to the 1950""s. The first publications on Hall thrusters appeared in the United States in the 1960""s, such as: G. R. Seikel and F. Reshotko, xe2x80x9cHall Current Ion Acceleratorxe2x80x9d, Bulletin of the American Physical Society, II (7) (1962) and C. O. Brown and E. A. Pinsley, xe2x80x9cFurther Experimental Investigations of Cesium Hall-Current Acceleratorxe2x80x9d, AIAA Journal, V.3, No 5, pp. 853-859, 1965.
Over the last thirty years, A. I. Morozov designed a series of high-efficiency Hall thrusters. See, for example, A. I. Morozov et al., xe2x80x9cEffect of the Magnetic field on a Closed-Electron-Drift Acceleratorxe2x80x9d, Sov. Phys. Tech. Phys. 17(3), pp. 482-487 (1972), A. I. Morosov, xe2x80x9cPhysical Principles of Cosmic Jet Propulsionxe2x80x9d, Atomizdat, Vol. 1, Moscow 1978, pp. 13-15, and A. I. Morozov and S. V. Lebedev, xe2x80x9cPlasma Opticsxe2x80x9d, in Reviews of Plasma Physics, Ed. by M. A. Leontovich, V.8, New York-London (1980).
H. R. Kaufman, xe2x80x9cTechnology of Closed Drift Thrustersxe2x80x9d, AIAA Journal Vol. 23 p. 71 (1983), reviews of the technology of Hall field thrusters, both in the context of other closed electron drift thrusters and in the context of other means of thrusting plasma. V. V. Zhurin et al., xe2x80x9cPhysics of Closed Drift Thrustersxe2x80x9d, Plasma Sources Science Technology Vol. 8, p. R1 (1999), further reviews the physics and more recent developments in the technology of Hall thrusters.
What remains a challenge is to develop a Hall thruster able to operate efficiently with minimal plume divergence. What is a further challenge is to accomplish such operation with the same thruster in several parameter regimes, such as at different input powers or at varying output thrusts. A number of issues arise with such variable operation of Hall current accelerators. These issues include decreased thruster efficiency for low mass flow rate and for low discharge voltages. At lower mass flow rates, lower atomic density in the channel results in an increased ionization mean free path of propellant atoms. A longer ionization length reduces the ionization efficiency and increases ion losses in the channel. Moreover, an extended ionization region produces a spread of ion energies, including slow ions. These slow ions are particularly vulnerable to radial accelerations and so contribute importantly to the plume divergence. This is a crucial issue even for non-variable operation. A similar effect would be incurred through the use of not easily ionized gases.
The present invention comprises an improvement over the prior art cited above by providing for efficient operation, with decreased plume divergence, and with capability for variable operation. The present invention discloses means of accomplishing these objectives through the placement of segmented electrodes along the inner and outer channel walls with the electrode segments held at specific potentials that lead to the improved operation.
The present invention comprises an improvement as well as over the following prior art:
U.S. Pat. No. 4,862,032 (xe2x80x9cEnd-Hall ion sourcexe2x80x9d, Kaufman et al., Aug. 29, 1989) discloses specifically that the magnetic field strength decreases in the direction from the anode to the cathode. The disclosure of the above referenced patent is hereby incorporated by reference.
Other design suggestions are disclosed in U.S. Pat. No. 5,218,271 (xe2x80x9cPlasma accelerator with closed electron driftxe2x80x9d, V. V. Egorov et al., Jun. 8, 1993) which contemplates a curved outlet passage. The disclosure of the above referenced patent is hereby incorporated by reference. U.S. Pat. No. 5,359,258 (xe2x80x9cPlasma accelerator with closed electron driftxe2x80x9d, Arkhipov et al., Oct. 25, 1994) contemplates improvements in magnetic source design by adding internal and external magnetic screens made of magnetic permeable material between the discharge chamber and the internal and external sources of magnetic field. The disclosure of the above referenced patent is hereby incorporated by reference.
U.S. Pat. No. 5,475,354 (xe2x80x9cPlasma accelerator of short length with closed electron driftxe2x80x9d, Valentian et al., Dec. 12, 1995) contemplates a multiplicity of magnetic sources producing a region of concave magnetic field near the acceleration zone in order better to focus the ions. The disclosure of the above referenced patent is hereby incorporated by reference. U.S. Pat. No. 5,581,155 (xe2x80x9cPlasma accelerator with closed electron driftxe2x80x9d, Morozov, et al., Dec. 3, 1996) similarly contemplates specific design optimizations of the conventional Hall thruster design, through specific design of the magnetic field and through the introduction of a buffer chamber. The disclosure of the above referenced patent is hereby incorporated by reference.
U.S. Pat. No. 5,763,989 (xe2x80x9cClosed drift ion source with improved magnetic fieldxe2x80x9d, H. R. Kaufman Jun. 9, 1998) contemplates the use of a magnetically permeable insert in the closed drift region together with an effectively single source of magnetic field to facilitate the generation of a well-defined and localized magnetic field, while, at the same time, permitting the placement of that magnetic field source at a location well removed from the hot discharge region. The disclosure of the above referenced patent is hereby incorporated by reference. U.S. Pat. No. 6,075,321 (xe2x80x9cHall field plasma accelerator with an inner and outer anodexe2x80x9d, V. J. Hruby, Jun. 13, 2000) contemplates an anode that can be part of either the inner or outer walls, rather than simply part of an inlet wall, but not a series of segmented electrodes for detailed control of the axial potential. The disclosure of the above referenced patent is hereby incorporated by reference.
U.S. Pat. No. 5,847,493 (xe2x80x9cHall effect plasma acceleratorxe2x80x9d, Yashnov et al., Dec. 8, 1998) proposes that the magnetic poles in an otherwise conventional Hall thruster be defined on bodies of material which are magnetically separate. The disclosure of the above referenced patent is hereby incorporated by reference.
U.S. Pat. No. 5,845,880 (xe2x80x9cHall effect plasma thrusterxe2x80x9d, Petrosov et al., Dec. 8, 1998) proposes a channel preferably flared outwardly at its open end so as to avoid erosion. The disclosure of the above referenced patent is hereby incorporated by reference.
The closest configuration in the literature to the present invention appears to be Russian Patent SU 1796777 A1 (Yu. M. Lisikov, V. V. Gopanchuk and I. B. Sorokin, xe2x80x9cStationary Plasma Thrusterxe2x80x9d, Applied Jun. 28, 1991, Issued: Feb. 23, 1993, Bulletin 7, in Russian). Lysikov et al. discloses an additional internal thermionic cathode, supplementary to the cathode compensator outside the acceleration region. The internal cathode is apparently placed where the magnetic field lines are approximately radial, which is approximately at the radial magnetic field maximum. The internal cathode is positioned on the discharge chamber apparently at the potential of the external cathode. In contrast to Lysikov et al., we disclose the design and use of emissive and non-emissive electrodes specifically configured so as to control and improve the voltage profile and thereby minimize the plume divergence. The disclosure of the above referenced patent is hereby incorporated by reference.
It is an object of this invention to provide an improved Hall plasma thruster by means of detailed control of the electric field.
It is a further object of this invention to provide an improved plasma thruster, which provides better focusing of the ion trajectories, thereby providing a more directional plume. A more tightly focused plasma plume reduces channel erosion, improves thrust, and facilitates integration with other satellite components.
The invention exploits the fact that the lines of magnetic force form surfaces of substantially constant electric potential. Since the magnetic field lines intersect the thruster channel, the potential distribution within the channel can be determined by imposing a potential distribution on the channel, through the placement of electrodes on the channel wall. The potential drop can then be imposed in a predetermined region of the thruster channel.
In the operation of a conventional Hall thruster, the total accelerating voltage, namely the voltage drop between the cathode and the anode, is fixed. However, the specific profile of the voltage drop between the anode and the cathode is dependent upon the details of the plasma flow and the magnetic field distribution. In order to control the electric potential in detail and, in particular, independent of the magnetic field, electrode segments are inserted along the plasma channel.
If the electrodes are not emissive, then an electrostatic plasma sheath will form in the vicinity of the electrode so as to shield the thruster interior from the electrode potential. This will generally be a deleterious effect, if not carefully designed, as ions will fall through a radial potential and strike the wall to balance the electron flux to the wall. However, if emissive electrode segments are employed, cold electrons are emitted from the wall, balancing the current of hot electrons to the wall, so that a radial sheath potential will not form. The ions are then not exposed to a radial potential drop. The ions then tend not to strike the wall and will produce a more tightly focused plasma plume. We disclose herein certain configurations of emissive and non-emissive electrodes to optimize thruster performance particularly by focusing the plume.
The present invention discloses an apparatus and method for thrusting plasma, utilizing a Hall thruster with segmented electrodes along the channel, which make the acceleration region as localized as possible. Also disclosed are methods of arranging the electrodes along the plasma channel so as to increase efficiency and minimize erosion and arcing. Also disclosed are methods of arranging the electrodes so as to produce a substantial reduction in plume divergence. The use of electrodes made of emissive material will reduce the radial potential drop within the channel, further decreasing the plume divergence. Also disclosed is a method of arranging and powering these electrodes so as to provide variable mode operation.
Since the magnetic field lines in a Hall thruster comprise magnetic surfaces at substantially the same electric potential, the voltage in the thruster interior may be substantially defined by imposing a specified electric potential on an electrode on the periphery of said interior region, such that the magnetic field line that permeates said interior thruster region also intersects said electrode. The method of specifying the potential on this field line is by inserting an electrode within the thruster channel, held at said potential, and such that said field line intersects said electrode.
This idea can be understood with reference to FIG. 1. FIG. 1 is a schematic representation of the plasma channel with segmented electrodes. Line 1Axe2x80x941A is a magnetic field line that extends from electrode segment 2 on channel wall 3 to an interior region in the thruster, which, which is approximately midway along the magnetic field line 1Axe2x80x941A. The magnetic field line extends to channel wall 5. Similarly, Line 1Bxe2x80x941B is a magnetic field line that extends from electrode segment 4 on channel wall 3 to an interior region in the thruster, which is approximately midway along the magnetic field line 1Bxe2x80x941B, and then similarly intersects the opposite channel wall 5. In a Hall thruster, lines 1Axe2x80x941A and 1Bxe2x80x941B would be substantially in the radial direction near the maximum of the magnetic field (see FIG. 2). For example, channel wall 3 could be the outer thruster wall and channel wall 5 could be the inner thruster wall, although the segmented electrodes could be placed against either or both walls so long as the same magnetic field lines is intersected by the electrode.
In the absence of plasma sheath effects, magnetic field line 1Axe2x80x941A tends to be at the same electric potential, since electrons can move freely along the field line to cancel any potential differences. Moreover, in a Hall thruster, electrons drift in the azimuthal direction, so that all field lines that intersect the channel at the same axial position tend to form surfaces of the same electric potential. The plasma sheath potential arises in order to balance the electron current to the channel wall by an ion current to the wall. If the electron axial flow is impeded by the magnetic field, then energetic electrons strike the wall faster than the ions do, until a sheath potential develops. However, if the electron temperature is small, or if the wall surface emits electrons, the sheath potential will be correspondingly small. The sheath potential impedes electrons from entering wall, but accelerates ions towards the wall. Accordingly, the sheath potential is a cause for ion plume divergence in Hall thrusters.
In one embodiment of the invention, the plasma sheath potential is small. Then all points along magnetic field line 1Axe2x80x941A are at approximately the same potential. Similarly, all points along magnetic field line 1Bxe2x80x941B are at approximately the same potential. The voltage source 6 establishes a potential drop between electrode segment 2 and electrode segment 4. Because each field line is substantially at the same potential along its own full length, said potential drop established between magnetic field line 1Axe2x80x941A and magnetic field line 1Bxe2x80x941B persists along the full length of both lines even throughout the thruster interior.
In a second embodiment of the invention, the plasma sheath effect may not be small. In said second embodiment, said electron potential along said magnetic field line 1Axe2x80x941A is determined partly by said potential imposed on electrode 2 and partly by electric sheath potential. However, by providing electrode 2 with emissive properties, said electrode 2 will emit electrons along magnetic field line 1Axe2x80x941A in such a manner as to cancel electric sheath potential.
In yet a third embodiment of the invention, the plasma sheath may not be small, yet electrodes without substantial emissive properties are employed. However, the electrodes are placed so as to minimize the plume divergence by providing for substantial axial accelerating potential in a precise and favorable region of the thruster channel.
FIG. 2 is a schematic representation Hall thruster with segmented electrode rings 7a and 7b on the outer ceramic channel wall 25. Line 0xe2x80x940 is an axis of symmetry. Segmented electrode 7b is near the thruster exit. Hollow cathode 8 emits electrons and neutralizes the flow of ions. An accelerating voltage drop is applied between anode 14 and hollow cathode 8, such that ions formed near the anode 14 are accelerated towards the thruster exit. The anode 14 can also be a as distributor. Magnetic field lines 10 extend from magnetic pole pieces 11 on the outer ceramic channel wall 25 and intersect magnetic pole pieces 12 on the inner ceramic channel wall 23. Electromagnetic coils 15 generate the magnetic field, which is guided through magnetic circuit 9 to the pole pieces. (An additional optional matched set of segmented electrodes 7c and 7d, placed on the inner channel wall 23 supplements said segmented electrode set 7a and 7b, such that said segmented electrode 7c intersects the same magnetic line of force as does electrode 7a and is held at the potential of electrode 7a. Similarly, said segmented electrode 7d intersects the same magnetic line of force as does electrode 7b and is held at the potential of electrode 7b.)
The electron current in the conventional Hall thruster provides the space charge neutralization and also assists in the ionization. While the current is primarily in the azimuthal direction, some axial current is necessary for the charge neutralization to occur. The electrons are normally introduced only through a cathode compensator, which could be a hollow cathode 8, outside of the main acceleration region of the ions and outside the region of intense magnetic fields. Thus, to neutralize the flow the electrons must travel axially towards the anode 14. Anode 14 can also serve as a gas distributor. Since power dissipated is proportional to current, the extent to which current is carried by these electrons is an unavoidable inefficiency. In addition to neutralizing the space charge within the acceleration region, the cathode compensator also serves to introduce electrons that neutralize the ion flow out of the thruster and eventually recombine with the ions. Thus the cathode compensator introduces electrons flowing in opposite axial directions: both electrons that flow back towards the anode and electrons that flow with the ion stream.
In a representative embodiment, the use of the set of segmented electrodes 7a and 7b is disclosed as an improvement. Rather than employing the cathode compensator outside the magnetic field for a dual purpose, we disclose how these functions may be separated. In the improved configuration, the cathode compensator outside the magnetic field need introduce only electrons that flow with the ions. The flow counter to the ions can be impeded by biasing the cathode 8 relative to segmented electrode 7b so that the ions experience a very small axial deceleration after leaving the acceleration region.
We disclose further that the set of segmented electrodes 7a and 7b provides a substantial voltage drop in a precise and predetermined location, thereby narrowing the ion plume and producing other advantages. We further disclose that said segmented electrodes 7a and 7b could be made of substantially emissive material not only to reduce deleterious effects of a plasma sheath but also to provide electrons necessary for ionization of the propellant gas. Each segment of an emissive electrode provides electrons by thermionic emission, secondary emission, field emission, capillary injection of electrons, or some other plasma producing means. The electrons so provided are available for charge neutralization, sheath reduction, or impact ionization of the neutral gas.
Note that the electrons need flow azimuthally only in the crossed electric and magnetic fields to provide the charge neutralization. Therefore, in yet another embodiment of the invention, a set of emissive segmented electrodes, such as set 7a and 7b, maintains the localized voltage drop at a precise and specified location within the acceleration region as well as providing for charge neutralization within the acceleration region. We disclose that the magnetic field that is imposed within the acceleration region may be allowed to be too large to permit electron axial current sufficient for ionization of the neutral gas. Instead, an additional emissive electrode segment, located between the acceleration region and the anode in a region of lower radial magnetic field, provides sufficient electrons for ionization of the neutral propellant gas. This additional electrode could be segmented electrode 7a, which is made sufficiently emissive not only to neutralize the electron sheath along the magnetic line of force intersecting it, but also to provide sufficient electrons for ionizing the upstream gas near the anode 14. In addition, a highly emissive segmented electrode 7b could provide sufficient electrons for neutralizing the accelerated ions, thus effectively serving also as the cathode-neutralizer 8.
Thus, it is a further object of the present invention to provide the electron current only where it is needed. The invention may thus be thought of also as a means of replacing some or all of the functions of the hollow cathode compensator 8. The consequence will be to reduce the electron power loss and thus improve the thruster efficiency of operation. Moreover, since the axial electron current is not essential in the acceleration region, higher magnetic fields can be used, without impeding the axial electron flow necessary for ionization. The use of high magnetic fields results in higher thrust density, since the thrust density cannot exceed the highest magnetic field energy density.